1. FIELD OF THE INVENTION
The present invention relates to space satellite system architecture, and more particularly, to a multi-mission spacecraft bus having improved structural, thermal and accessibility characteristics.
2. DESCRIPTION OF RELATED ART
Space satellite systems play a critical role in today's modern society. These orbiting satellites perform a diverse number of military and civilian missions, including earth mapping, navigation, communications, atmospheric/environmental sensing, early warning/ targeting, and scientific research and experimentation. Earth orbiting satellites can either be positioned into a high earth orbit, or geosynchronous orbit, in which the satellite appears to hover over a specific equatorial point of the earth, or low earth orbit, in which the satellite rotates many times around the earth in a given day, obtaining data on the entire globe. At any given moment, hundreds of space satellites are in orbit, providing dedicated and specific services to its ground users.
Although many such satellites are built and flown each year, their total numbers are still very low. For this reason, many conventional manufacturing techniques, such as common to the building of automobiles or television sets, have failed to be incorporated into the development of spacecraft systems. This is because the typical satellite system is designed and built for a specific intended purpose, and most of its components are custom made for that purpose. Interchangeability of components between different satellite system is extremely rare. Purchasers of space satellite systems usually only buy a small number at a time and the manufacturers have little incentive to introduce mass production techniques. One drawback of this is that as new ideas or technologies are developed, it is virtually impossible to incorporate these new techniques into existing designs, short of redesigning the entire spacecraft system.
It has long been desired to provide a basic multi-mission spacecraft structure which can be adapted for use in a variety of roles. This basic structure, known as a spacecraft bus, would interface with the launch vehicle, and provide basic housekeeping functions for the specific payload. The payload is the specific sensor or instrument which provides the core function of the specific satellite mission. The housekeeping functions include power generation, station keeping, thermal control, and ground communications. To provide power generation, many satellites use solar panels. The solar panels have light sensitive cells which convert light energy into electricity. The solar panels must be continually adjusted to maintain an optimum angle with the sun. To provide thermal control, a spacecraft system must be able to eliminate excess heat. Radiators are often used to eliminate such heat, and thermal blankets are used to insulate sensitive components and reflect light energy away from the spacecraft.
Station keeping involves the maintenance of the spacecraft in its proper orbital position. Spacecraft systems are typically either spin stabilized, or three access stabilized. A spin stabilized spacecraft continually rotates upon a center axis, behaving like a gyroscope. This gyroscopic action tends to maintain the spacecraft pointing in a given direction. Minor adjustments in the direction can be achieved through the use of thrusters which expel gas in a given direction to reorient the position of the vehicle. Three access stabilized satellites do not spin, but instead rely upon a complex network of sensors and positioning systems. The sensors include gyroscopes which can provide instantaneous position information. This information can be provided to the positioning system which compensates the actual spacecraft position to return it to its intended position. These positioning systems often include thrusters, as discussed above, as well as momentum wheels and reaction wheels. Both types of wheels operate on the principal that their spinning can either impart or eliminate satellite motion. The momentum wheels can be spun by the use of a motor, and their spinning can cause the satellite to rotate in the opposite direction. The reaction wheels are free to move in response to other forces acting upon the satellite to absorb the undesired forces. Stored energy, or rotational momentum, built up in the momentum or reaction wheels can be dissipated through the controlled usage of the thrusters.
Spacecraft ground communication is necessary to operate the spacecraft, and includes an information receiving and transmitting capability. An onboard transmitter/receiver is provided to decode signals received from the ground and relay them either to the payload or to the intended housekeeping system. Antennas are typically provided on an outer portion of the spacecraft structure to receive and transmit these signals. Frequently the antenna must be gimballed in order to maintain an optimal pointing angle with the associated ground transmitting system.
Each of these diverse systems must be housed together in a compact bus structure. The bus must also be capable of securing to the launching system, and be able to separate itself from the launching system. Moreover, the bus must be capable of withstanding the enormous loading which occurs during the launch and separation.
A typical prior art bus configuration is shown generally at 10 in FIG. 1. This configuration, known as a "thrust tube," has a generally cylindrical structure 16, having upper and lower receiving flanges shown at 14 and 12, for mounting to the launch vehicle and to the payload, respectively. The thrust tube design approach has many desirable structural features. Structural load paths are concurrent with the launch vehicle and transition directly from the launch vehicle separation system and up through the payload interface. A propulsion tank 18 can be typically provided at the center of the thrust tube structure 16. The thrust tube shape provides a large, closed section structure having efficient shear, torsional, and longitudinal load carrying capability. A non-load bearing frame 24 can mount to the structure 16, which supports external equipment, such as solar panels 22. The compact and symmetric shape of the thrust tube design makes it an ideal candidate for use in a spin stabilized configuration.
However, the structural efficiency of the thrust tube is balanced by the increased manufacturing complexity and inefficient component integration. It is relatively difficult to manufacture a balanced, circular structure, and to custom fit the components which must mount to the structure. Essentially, spacecraft designers must put a square peg (internal electronic components) into a round hole (thrust tube structure). Moreover, once the components are custom fit into the internal portions of the thrust tube, their subsequent removal requires a disassembly of the entire bus structure 16. Such removal may be necessary if internal components fail during ground test, or if technological advancements are desired to upgrade the components.
An alternative to the thrust tube structure is known as the outside shell structure, which is shown generally at 30 in FIG. 2. The outside shell structure is a variation of the thrust tube structure, being generally circular but formed of a plurality of flat surfaces. The outside shell structure shown in FIG. 2 is octagonally shaped, having eight external panels 32 which are individually removable. The top 37 and bottom 39 of the shell can be sealed, with the bottom portion providing an adaptor interface with the launch vehicle. The upper surface 37 provides a mounting point for the payload. As with the thrust tube design, the outside shell alternative uses the external panels 32 to carry the primary structural loads. The interior volume can accommodate the propulsion tanks 34 as well as the electronic equipment. The multi-sided structure creates pie-shaped equipment mounting areas, which after incorporation of the propulsive elements limits the available equipment stowage volume. This limited electronics area constrains the ability of the design to incorporate developmental equipment, as with the thrust tube. An additional problem with the outside shell design is that the external panels used for internally mounting the equipment are necessary for the vehicle structure. The equipment panels cannot be removed, nor can the interior equipment be accessed, without removal of these primary structural elements.
An additional problem with both the thrust tube and the outside shell designs, is that of effective thermal management. In each design, the onboard electronics equipment is stowed within the large, central compartment formed by the tube or shell. Heat generated by these electronic systems must be removed from the structure to prevent damage due to overheating the delicate components. Heat generated within the structure would rapidly flow to each of the other components within the compartment. Thus, the shape of the structures renders it difficult to thermally isolate critical components and reject the built-up heat.
Thus, it would be desirable to provide a standard multi-mission spacecraft bus structure having improved load carrying and equipment mounting capability over the prior art. It would also be desirable to provide a multi-mission bus structure having components which are accessible from external to the structure. It would be still further desirable to provide a multi-mission bus structure having decentralized thermal removal capabilities.